Sectioned engine structure for a gas turbine engine

ABSTRACT

An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a stationary engine structure. The stationary engine structure includes a diffuser, a combustor, an engine case and a plenum. The combustor is disposed within the plenum. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. A first section of the stationary engine structure is formed as a first monolithic body. The first section includes the diffuser and the combustor. A second section of the stationary structure is formed as a second monolithic body. The second section is configured as or otherwise includes the engine case.

BACKGROUND OF THE DISCLOSURE 1. Technical Field

This disclosure relates generally to a gas turbine engine and, moreparticularly, to a stationary structure for the gas turbine engine.

2. Background Information

A gas turbine engine includes a stationary engine structure for housingand/or supporting internal rotating components of the gas turbineengine. A typical stationary engine structure includes a plurality oftubular axial case segments. These tubular axial case segments arearranged sequentially along an axial centerline of the gas turbineengine and axially connected together by flange connections. While sucha stationary engine structure has various benefits, there is still roomin the art for improvement. There is a need in the art therefore for animproved stationary engine structure as well as methods formanufacturing and assembling stationary engine structure components.

SUMMARY OF THE DISCLOSURE

According to an aspect of the present disclosure, an assembly isprovided for a gas turbine engine. This gas turbine engine assemblyincludes a stationary engine structure. The stationary engine structureincludes a diffuser, a combustor, an engine case and a plenum. Thecombustor is disposed within the plenum. The engine case forms aperipheral boundary of the plenum. A gas path extends sequentiallythrough the diffuser, the plenum and the combustor. A first section ofthe stationary engine structure is formed as a first monolithic body.The first section includes the diffuser and the combustor. A secondsection of the stationary structure is formed as a second monolithicbody. The second section is configured as or otherwise includes theengine case.

According to another aspect of the present disclosure, another assemblyis provided for a gas turbine engine. This gas turbine engine assemblyincludes a diffuser, a combustor, a duct wall and an engine wall. Thediffuser includes an inner diffuser wall and an outer diffuser wall. Thecombustor includes an inner combustor wall, an outer combustor wall anda bulkhead extending between and connected to the inner combustor walland the outer combustor wall. The engine wall includes a side wall andan end wall. The side wall projects out from the end wall to the outerdiffuser wall. The side wall is brazed to the outer diffuser wall. Theend wall projects out from the side wall to the duct wall. The end wallis brazed to the duct wall.

According to still another aspect of the present disclosure, amanufacturing method is provided. During this method, an enginestructure preform is formed. The engine structure preform includes afirst section preform and a second section preform formed integral withthe first section preform. The first section preform includes a diffuserand a combustor. The second section preform is configured as orotherwise includes an engine case. The first section preform isseparated from the second section preform to respectively provide afirst section and a second section that is discrete from the firstsection. The second section is attached to the first section to providean engine structure. The combustor is disposed within a plenum of theengine structure. The engine case forms a peripheral boundary of theplenum. A gas path extends sequentially through the diffuser, the plenumand the combustor.

The forming may include additive manufacturing the engine structurepreform. In addition or alternatively, the attaching may include bondingthe second section to the first section.

The assembly may include a monolithic body including the diffuser, thecombustor and the duct wall.

The diffuser may include a first wall, a second wall and a plurality ofvanes. Each of the vanes may be within the gas path and may extendbetween the first wall and the second wall.

The combustor may be configured as a reverse-flow combustor.

The stationary engine structure may also include a turbine nozzledownstream of the combustor along the gas path. The first section mayinclude the turbine nozzle.

The diffuser and the turbine nozzle may share a common wall.

The turbine nozzle may include a first wall, a second wall and aplurality of vanes. Each of the vanes may be within the gas path and mayextend between the first wall and the second wall.

The assembly may also include a turbine rotor. The stationary enginestructure may also include a turbine case housing the turbine rotor andforming a peripheral boundary of the gas path. The first section mayalso include the turbine case.

The stationary engine structure may also include an exhaust duct forminga peripheral boundary of the gas path. The first section may alsoinclude the exhaust duct.

The second section may circumscribe the first section.

The second section may be bonded to the first section through a buttjoint.

The second section may be bonded to the first section through a splicejoint.

The engine case may extend axially along and circumferentially about anaxis. The engine case may include a side wall and an end wall. The sidewall may project axially out from the end wall to an axial end of theengine case. The engine case may be attached to the first section at theaxial end. The end wall may project radially in from the side wall to aradial end of the engine case. The engine case may be attached to thefirst section at the radial end.

The stationary engine structure may also include a fuel conduit and anozzle. The nozzle may be configured to receive fuel from the fuelconduit and inject the fuel into a volume of the combustor. The secondsection may also include the fuel conduit and/or the nozzle.

The stationary structure may also include a fuel manifold outside of theengine case. The fuel manifold may be configured to supply the fuel tothe fuel conduit. The second section may also include the fuel manifold.

At least a portion of the fuel conduit may project out from the enginecase into the plenum towards the fuel injector.

The stationary engine structure may also include an inlet section and acompressor case. The compressor case may form a peripheral boundary ofthe gas path between the inlet section and the diffuser. A third sectionof the stationary structure may include the inlet section and thecompressor case. The third section may be attached to the first section.

The present disclosure may include any one or more of the individualfeatures disclosed above and/or below alone or in any combinationthereof.

The foregoing features and the operation of the invention will becomemore apparent in light of the following description and the accompanyingdrawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic side sectional illustration of a gas turbineengine.

FIG. 2 is an aft end view illustration of a stationary engine structurefor the gas turbine engine.

FIG. 3 is a schematic side sectional illustration of a central portionof the gas turbine engine.

FIG. 4 is a schematic perspective illustration of a portion of a fueldelivery system.

FIG. 5 is a partial side sectional illustration of an outer section ofthe stationary engine structure bonded to an inner section of thestationary engine structure.

FIG. 6 is a schematic illustration of a manufacturing system.

FIG. 7 is a flow diagram of a method for forming a gas turbine engine.

FIGS. 8A and 8B are schematic illustrations of various preforms of thestationary engine structure.

DETAILED DESCRIPTION

FIG. 1 is a side sectional illustration of a gas turbine engine 20. Thegas turbine engine 20 of FIG. 1 is configured as a single spool,radial-flow turbojet turbine engine. This gas turbine engine 20 isconfigured for propelling an aircraft such as, but not limited to, anunmanned aerial vehicle (UAV), a drone or any other manned or unmannedaircraft or self-propelled projectile. The present disclosure, however,is not limited to such an exemplary turbojet turbine engineconfiguration nor to an aircraft propulsion system application. Forexample, the gas turbine engine 20 may alternatively be configured as anauxiliary power unit (APU) or an industrial gas turbine engine.

The gas turbine engine 20 of FIG. 1 extends axially along an axialcenterline 22 between a forward, upstream engine inlet 24 and an aft,downstream engine exhaust 26. This axial centerline 22 may also be arotational axis for various components within the gas turbine engine 20.

The gas turbine engine 20 includes a compressor section 28, a combustorsection 30 and a turbine section 32. The gas turbine engine 20 alsoincludes a stationary engine structure 34. This stationary enginestructure 34 houses the compressor section 28, the combustor section 30and the turbine section 32. The stationary engine structure 34 of FIG. 1also forms an inlet section 36 and an exhaust section 38 for the gasturbine engine 20, where the inlet section 36 forms the engine inlet 24and the exhaust section 38 forms the engine exhaust 26.

The engine sections 36, 28, 30, 32 and 38 are arranged sequentiallyalong a core gas path 40 that extends through the gas turbine engine 20from the engine inlet 24 to the engine exhaust 26. Each of the enginesections 28 and 32 includes a respective rotor 42, 44. Each of theserotors 42, 44 includes a plurality of rotor blades arrangedcircumferentially around and connected to at least one respective rotordisk. The rotor blades, for example, may be formed integral with ormechanically fastened, welded, brazed, adhered and/or otherwise attachedto the respective rotor disk(s).

The compressor rotor 42 may be configured as a radial flow rotor. Theturbine rotor 44 may also or alternatively be configured as a radialflow rotor. The compressor rotor 42 is connected to the turbine rotor 44through an engine shaft 46. This shaft 46 is rotatably supported by thestationary engine structure 34 through a plurality of bearings 48A and48B (generally referred to as 48); e.g., rolling element bearings,journal bearings, etc.

The combustor section 30 includes an annular combustor 50 with anannular combustion chamber 52. The combustor 50 of FIG. 1 is configuredas a reverse flow combustor. Inlets ports 54 (e.g., dilution chutes)into the combustion chamber 52, for example, may be arranged at (e.g.,on, adjacent or proximate) and/or towards an aft bulkhead wall 56 (e.g.,bulkhead, dome, etc.) of the combustor 50. An outlet from the combustor50 may be arranged axially aft of an inlet to the turbine section 32.The combustor 50 may also be arranged radially outboard of and/oraxially overlap at least a (e.g., aft) portion of the turbine section32. With this arrangement, the core gas path 40 of FIG. 1 reverses itsdirections (e.g., from a forward-to-aft direction to an aft-to-forwarddirection) a first time as the gas path 40 extends from a combustorplenum 58 surrounding the combustor 50 into the combustion chamber 52.The core gas path 40 of FIG. 1 then reverses its direction (e.g., fromthe aft-to-forward direction to the forward-to-aft direction) a secondtime as the gas path 40 extends from the combustion chamber 52 into theturbine section 32.

During operation, air enters the gas turbine engine 20 through the inletsection 36 and its engine inlet 24. The inlet section 36 directs thisair from the engine inlet 24 into the core gas path 40 and thecompressor section 28. The engine inlet 24 of FIG. 1 thereby forms aforward, upstream inlet to the core gas path 40 and the compressorsection 28. The air within the core gas path 40 may be referred to ascore air.

The core air is compressed by the compressor rotor 42 and directedthrough an annular diffuser 60 and the plenum 58 into the combustionchamber 52. Fuel is injected and mixed with the compressed core air toprovide a fuel-air mixture. This fuel-air mixture is ignited within thecombustion chamber 52, and combustion products thereof flow through theturbine section 32 and cause the turbine rotor 44 to rotate. Thisrotation of the turbine rotor 44 drives rotation of the compressor rotor42 and, thus, compression of the air received from the engine inlet 24.The exhaust section 38 receives the combustion products from the turbinesection 32. The exhaust section 38 directs the received combustionproducts out of the gas turbine engine 20 to provide forward enginethrust.

The stationary engine structure 34 of FIG. 1 may include some or allstationary engine components included in the gas turbine engine 20.Herein, the term stationary may describe a component that does notrotate with the rotating assembly (e.g., an assembly of the rotors 42and 44 and the shaft 46) during gas turbine engine operation. Astationary component, for example, may refer to any component thatremains stationary during gas turbine engine operation such as, but notlimited to, a wall, a liner, a strut, a fixed vane, a fuel nozzle, aconduit, etc.

The stationary engine structure 34 of FIG. 1 is configured as agenerally tubular structure. The stationary engine structure 34, forexample, extends axially along the axial centerline 22 from the inletsection 36 to the engine section 38. The stationary engine structure 34extends circumferentially about (e.g., completely around) the axialcenterline 22 such that the stationary engine structure 34 has, forexample, a full hoop geometry; see also FIG. 2 .

The stationary engine structure 34 includes one or more case walls. Thestationary engine structure 34 of FIG. 3 , for example, includes anouter compressor wall 62, an outer diffuser wall 64 of the diffuser 60,an inner diffuser wall 66 of the diffuser 60, a plenum side (e.g.,outer) wall 68, a plenum end wall 70, an outer combustor wall 72 of thecombustor 50, an inner combustor wall 74 of the combustor 50, thebulkhead wall 56 of the combustor 50, an inner turbine wall 76 of aturbine duct 78, and an exhaust wall 80 of an exhaust duct 82 (see alsoFIG. 1 ). At least a portion or an entirety of each of the case walls62, 64, 66, 68, 72, 74, 76 and/or 80 of FIG. 3 , for example, isgenerally tubular. At least a portion or an entirety of each of the casewalls 56 and/or 70 of FIG. 3 is generally annular.

The compressor wall 62 extends axially along the axial centerline 22between and is connected to the inlet section 36 and the outer diffuserwall 64. The compressor wall 62 of FIG. 3 circumscribes, axiallyoverlaps and thereby houses the compressor rotor 42.

The outer diffuser wall 64 extends axially along the axial centerline 22between and is connected to the compressor wall 62 and the plenum sidewall 68. The outer diffuser wall 64 is spaced radially outboard from,axially overlaps and circumscribes the inner diffuser wall 66. The outerdiffuser wall 64 of FIG. 3 thereby forms an outer peripheral boundary ofthe core gas path 40 through the diffuser 60.

The inner diffuser wall 66 may be connected to outer combustor wall 72.The inner diffuser wall 66 of FIG. 3 , for example, projects axially outfrom the outer combustor wall 72 and extends axially towards (e.g., to)an aft, downstream end of an inner platform of the compressor rotor 42.This inner diffuser wall 66 forms an inner peripheral boundary of thecore gas path 40 within the diffuser 60. The inner diffuser wall 66 mayalso be configured as an outer turbine wall. The inner diffuser wall 66of FIG. 3 , for example, may also form an outer peripheral boundary ofthe core gas path 40 within a (e.g., upstream) portion of the turbinesection 32. The inner diffuser wall 66 of FIG. 3 circumscribes, axiallyoverlaps and may thereby house a (e.g., upstream) portion of the turbinerotor 44.

The plenum side wall 68 extends axially along the axial centerline 22between and is connected to the outer diffuser wall 64 and the plenumend wall 70. The plenum side wall 68 of FIG. 3 circumscribes, axiallyoverlaps and thereby houses the combustor 50 and its outer combustorwall 72. The plenum side wall 68 is radially spaced outward from thecombustor 50 and its outer combustor wall 72. The plenum side wall 68forms an outer peripheral boundary of the plenum 58.

The plenum end wall 70 extends radially (and axially along the axialcenterline 22) between and is connected to the plenum side wall 68 andthe exhaust wall 80. The plenum end wall 70 is axially spaced from thecombustor 50 and its bulkhead wall 56. The plenum end wall 70 forms anaxial end peripheral boundary of the plenum 58.

The outer combustor wall 72 extends axially along the axial centerline22 between and may be connected to the bulkhead wall 56 and the innerdiffuser wall 66. More particularly, the outer combustor wall 72 extendsaxially to and may be connected to an outer platform 84 of a turbinenozzle 86; e.g., an exit nozzle from the combustion chamber 52. Thisnozzle outer platform 84 of FIG. 3 is configured as part of the innerdiffuser wall 66; however, the walls 66 and 84 may be discrete from oneanother in alternative embodiments.

The inner combustor wall 74 is connected to the bulkhead wall 56. Thisinner combustor wall 74 projects axially along the axial centerline 22out from the bulkhead wall 56 towards the turbine nozzle 86 and itsinner platform 88. This nozzle inner platform 88 of FIG. 3 is configuredas part of the inner turbine wall 76; however, the walls 76 and 88 maybe discrete from one another in alternative embodiments.

The bulkhead wall 56 extends radially between the outer combustor wall72 and the inner combustor wall 74. The bulkhead wall 56 is connected toan aft end portion of the outer combustor wall 72 and an aft end portionof the inner combustor wall 74. With this arrangement, the combustorcase walls 56, 72 and 74 collectively form peripheral boundaries of thecombustion chamber 52 within the combustor 50.

The inner turbine wall 76 may be wrapped around a downstream end portionof the inner combustor wall 74. An upstream portion of the inner turbinewall 76 of FIG. 3 (e.g., the inner platform 88), for example,circumscribes and axially overlaps the downstream end portion of theinner combustor wall 74. This upstream portion extends axially along theaxial centerline 22 (in the aft-to-forward direction) to a turningportion of the inner turbine wall 76. A downstream portion of the innerturbine wall 76 projects axially (in the forward-to-aft direction) awayfrom the inner turbine wall turning portion to the exhaust wall 80. Theinner turbine wall 76 is circumscribed and axially overlapped by thecombustor 50 and its inner combustor wall 74. The inner turbine wall 76is also spaced radially inboard from the combustor 50 and its innercombustor wall 74. The inner turbine wall 76 of FIG. 3 forms an innerperipheral boundary of the plenum 58, where the combustor 50 is disposedwithin and is substantially surrounded by the plenum 58. The innerturbine wall 76 forms an outer peripheral boundary of the core gas path40 within a (e.g., downstream) portion of the turbine section 32. Theinner turbine wall 76 of FIG. 3 also circumscribes, axially overlaps andthereby houses a (e.g., downstream) portion of the turbine rotor 44.

The exhaust wall 80 is connected to the inner turbine wall 76. Theexhaust wall 80 of FIG. 1 projects axially out from the inner turbinewall 76 to the aft engine exhaust 26.

The stationary engine structure 34 may include one or more internalsupport structures with one or more support members. Examples of thesupport members include, but are not limited to, struts, structuralguide vanes, bearing supports, bearing compartment walls, etc. Thestationary engine structure 34 of FIG. 3 , for example, includes aforward support structure 90 to support the forward bearing 48A and anaft support structure 92 to support the aft bearing 48B. The stationaryengine structure 34 of FIG. 3 also includes an inlet nozzle 94, adiffuser nozzle 96 and the turbine nozzle 86.

The inlet nozzle 94 may be configured to condition the core air enteringthe compressor section 28. The inlet nozzle 94 of FIG. 3 , for example,includes one or more inlet guide vanes 98 configured to impart swirl tothe core air. These inlet guide vanes 98 are arranged circumferentiallyabout the axial centerline 22 in an annular array. Each of the inletguide vanes 98 extends radially across the gas path 40.

The diffuser nozzle 96 may be configured to condition the core airleaving the compressor section 28 and entering the plenum 58. Thediffuser nozzle 96 of FIG. 3 , for example, includes one or morediffuser guide vanes 100 configured to impart swirl to the core air.These diffuser guide vanes 100 are arranged circumferentially about theaxial centerline 22 in an annular array. Each of the diffuser guidevanes 100 extends radially across the gas path 40. More particularly,each of the diffuser guide vanes 100 extends radially between and isconnected to the inner diffuser wall 66 and the outer diffuser wall 64.

The turbine nozzle 86 may be configured to condition the combustionproducts exiting the combustor 50 and its combustion chamber 52. Theturbine nozzle 86 of FIG. 3 , for example, includes one or more turbineguide vanes 102 configured to impart swirl to the combustion products.These turbine guide vanes 102 are arranged circumferentially about theaxial centerline 22 in an annular array. Each of the turbine guide vanes102 extends radially across the gas path 40. More particularly, each ofthe turbine guide vanes 102 extends radially between and is connected tothe turbine nozzle outer and inner platforms 84 and 88.

Referring to FIG. 4 , the stationary engine structure 34 may alsoinclude one or more components of a fuel delivery system. These fueldelivery system components may include a fuel feed line 104 (e.g., aninlet conduit), a fuel manifold 106, one or more fuel conduits 108 andone or more fuel injectors 110. The fuel manifold 106 is configured tosupply fuel received from a fuel source (e.g., a fuel reservoir, a fuelpump, etc.) through the fuel feed line 104 to the fuel conduits 108,which fuel conduits 108 may be fluidly coupled to the fuel manifold 106in parallel. Referring to FIG. 5 , each of the fuel conduits 108 isconfigured to direct the fuel received from the fuel manifold 106 to arespective one of the fuel injectors 110. Each of the fuel injectors 110is configured to direct (e.g., inject) the fuel received from therespective fuel conduit 108 into a respective volume of the combustor 50for subsequent combustion within the combustion chamber 52. Thecombustor volume may be a respective one of the ports 54 in a sidewallof the combustor 50 (e.g., the outer combustor wall 72), where the port54 may be configured as a dilution chute, a flow guide, an orifice orany other opening and/or passage through the combustor sidewall and intothe combustion chamber 52.

The fuel manifold 106 of FIG. 4 extends circumferentially about (e.g.,completely around) the axial centerline 22. Referring to FIG. 5 , thefuel manifold 106 is arranged at an exterior of an outer engine case 112(e.g., a plenum case, a combustor section case, etc.) of the stationaryengine structure 34, which outer engine case 112 may include the plenumside wall 68 and the plenum end wall 70. An entirety of the fuelmanifold 106, for example, may be located outside of the plenum 58 and,more generally, outside of the engine case 112 and its plenum end wall70. The fuel manifold 106 is connected to and extends longitudinally(e.g., in a circumferential direction; see FIG. 4 ) along the enginecase 112.

The fuel conduits 108 of FIG. 4 are distributed circumferentially aboutthe axial centerline 22 in an annular array. Each of these fuel conduits108 extends longitudinally between and is fluidly coupled to the fuelmanifold 106 and a respective one of the fuel injectors 110. Referringto FIG. 5 , each fuel conduit 108 is arranged at an interior of theengine case 112. Each fuel conduit 108 of FIG. 5 , for example, isconnected to the engine case 112 and its plenum end wall 70. Each fuelconduit 108 projects (e.g., axially) out from the engine case 112 andits plenum end wall 70 into the plenum 58 towards (e.g., to) therespective fuel injector 110. A first (e.g., upstream) segment 114 ofeach fuel conduit 108 of FIG. 5 extends axially along and may beintegral with the engine case 112 and one or more of its walls 68 and/or70; e.g., the plenum end wall 70. A second (e.g., downstream) segment116 of each fuel conduit 108 of FIG. 5 projects axially out from theconduit first segment 114, away from the engine case 112, towards (e.g.,to) the respective fuel injector 110. This conduit second segment 116 isradially spaced (e.g., separated) from the engine case 112 and one ormore of its walls 68 and/or 70; e.g., the plenum side wall 68. A distalend of each fuel conduit 108 and/or each fuel injector 110, however, maybe structurally supported within the plenum 58 by a support 118; e.g., astrut, a vane, a post, a beam, etc. This support 118 is configured tomaintain a position of the fuel injector 110 during gas turbine engineoperation. Without the support 118, for example, thermal expansionand/or contraction of the fuel conduit 108 may lead to (e.g., axialand/or circumferential) misalignment of the respective fuel injector 110and port 54; e.g., dilution chute. Of course, in other embodiments, anentirety of one or more or all of the fuel conduits 108 may each beformed integral with the engine case 112 to obviate the need for therespective support 118.

The fuel injectors 110 of FIG. 4 are distributed circumferentially aboutthe axial centerline 22 in an annular array. Each of these fuelinjectors 110 is fluidly coupled with a respective one of the fuelconduits 108. Referring to FIG. 5 , each of the fuel injectors 110 is(e.g., circumferentially and/or axially) aligned with and located nextto the respective port 54; e.g., dilution chute. Each fuel injector 110is disposed at a distal end of the respective fuel conduit 108, and isformed integral with the respective fuel conduit 108 and support 118.Each fuel injector 110, for example, may be configured as a nozzle(e.g., an outlet orifice) in a sidewall of the respective fuel conduit108. Alternatively, one or more of the fuel injectors 110 may be formeddiscrete from the respective fuel conduit 108.

The stationary engine structure 34 of FIG. 3 is configured from aplurality of discrete sections of the engine structure 34; e.g., enginesub-structures. The stationary engine structure 34 of FIG. 3 , forexample, at least or only includes an upstream section 120, an innerdownstream section 121 and an outer downstream section 122. The upstreamsection 120 may include an entirety (or at least a portion) of any oneor more or all of the engine structure elements 36, 62, 90, 94 and 98.The inner downstream section 121 may include an entirety (or at least aportion) of any one or more or all of the engine structure elements 50,56, 64, 66, 72, 74, 76, 78, 80, 82, 84, 86, 88, 92, 96, 100, 102; seealso FIG. 1 . The outer downstream section 122 may include an entirety(or at least a portion) of any one or more or all of the enginestructure elements 68, 70, 104 (see FIGS. 2 and 4 ), 106, 108, 110, 112,114, 116, 118 and 134. With such an arrangement, the stationary enginestructure 34 may be configured from (include) relatively few discreteparts (e.g., discretely formed bodies) while still facilitatinginspection and/or finishing (e.g., machining, coating, etc.) interiorsurfaces of the stationary engine structure 34. For example, forming thedownstream sections 121 and 122 discrete from one another facilitatesinspection and/or finishing of the fuel injectors 110 and theirorifices, inspection and/or finishing of the ports 54 (e.g., dilutionchutes), etc.

Each of the engine structure sections 120-122 may be formed as amonolithic body. Herein, the term monolithic may described an apparatuswhich is formed as a single unitary body. Each engine structure section120, 121, 122, for example, may be additively manufactured, cast,machined and/or otherwise formed as an integral, unitary body. Bycontrast, a non-monolithic body may include parts that are discretelyformed from one another, where those parts are subsequently mechanicallyfastened and/or otherwise attached to one another.

The upstream section 120 is mated with and connected to a forward,upstream end of the inner downstream section 121. The upstream section120 of FIG. 3 , for example, is attached to the inner downstream section121 via at least one mechanical joint 124; e.g., a bolted flangeconnection. However, in other embodiments, the upstream section 120 mayalso or alternatively be attached to the inner downstream section 121via at least one bonded joint; e.g., a brazed connection, a weldedconnection, etc.

The outer downstream section 122 is mated with and connected to theinner downstream section 121. The inner downstream section 121 of FIG. 5, for example, is received within a bore of the outer downstream section122, where the outer downstream section 122 and its engine case 112circumscribe and axially overlap the combustor 50 and the exhaust duct82. The outer downstream section 122 is connected to the innerdownstream section 121 at a forward axial end 126 of the engine case 112and a radial inner end 128 of the engine case 112. The plenum side wall68 of FIG. 5 , for example, is brazed, welded and/or otherwise bonded tothe outer diffuser wall 64 at an (e.g., annular) axial interface 130;e.g., a butt joint. An annular end (e.g., edge) of the plenum side wall68 of FIG. 5 , for example, is aligned with and axially engages (e.g.,via bonding material such as braze material) an annular end (e.g., edge)of the outer diffuser wall 64. The plenum end wall 70 of FIG. 5 isbrazed, welded and/or otherwise bonded to the exhaust duct 82 and itsexhaust wall 80 (and/or the turbine duct 78 and its inner turbine wall76) at a (e.g., tubular) radial interface 132; e.g., a splice joint suchas an overlap joint. The engine case 112 of FIG. 5 , for example,includes a tubular flange 134 that projects axially (e.g., in an aft,downstream direction) out from the plenum end wall 70 at the radialinner end 128 to an axial distal end. An inner surface of this tubularflange 134 circumscribes, axially overlaps and radially engages (e.g.,via bonding material such as braze material) an outer surface of theexhaust duct 82 and its exhaust wall 80 (and/or the turbine duct 78 andits inner turbine wall 76).

The axial interface 130 and/or the radial interface 132 may bepositioned relatively far from one or more of the fuel delivery systemcomponents. The axial interface 130 of FIG. 5 , for example, is locatedat (e.g., on, adjacent or proximate) the diffuser 60 and an outlet fromthe combustor 50. The radial interface 132 of FIG. 5 is located at(e.g., on, adjacent or proximate) the exhaust duct 82 (and/or theturbine duct 78) and a radial inner side of the combustor 50. With thisarrangement, heat transferred into the outer downstream section 122 andits engine case 112 during bonding of the engine structure sections 121and 122 may significantly dissipate prior to reaching the fuel deliverysystem components 108 and/or 110. The fuel delivery system components108 and/or 110 may therefore be subject to very little or no thermaldistortion during the bonding of the engine structure sections 121 and122. The fuel injectors 110 may thereby remain properly positionedrelative to (e.g., aligned with) the respective ports 54. By contrast,if the interfaces 130 and 132 are positioned too close to (e.g., nextto) the fuel delivery system components 108 and/or 110, the fueldelivery system components 108 and/or 110 may be subject to thermaldistortion during bonding and the fuel injectors 110 may becomemisaligned with (e.g., offset from) the respective ports 54.

FIG. 6 illustrates a system 136 for manufacturing a stationary enginestructure such as, but not limited to, the stationary engine structure34 of FIG. 1 . This manufacturing system 136 (e.g., an additivemanufacturing system) includes a build chamber 138 defining a buildspace 140 for manufacturing the stationary engine structure 34 (or oneor more of its components). The manufacturing system 136 also includesan additive manufacturing apparatus 142; e.g., a laser powder bed fusion(LPBF) apparatus.

The additive manufacturing apparatus 142 is configured to build thestationary engine structure 34 (or one or more of its components) or apreform thereof within the build space 140 in a layer-by-layer fashion.For example, the additive manufacturing apparatus 142 may deposit afirst layer 144A of powder over a support surface 146 within the buildspace 140. The additive manufacturing apparatus 142 may thereafterselectively solidify (e.g., sinter) a select portion of the first layer144A of powder to form a first portion 148A (e.g., layer, slice) of thestationary engine structure 34 (or one or more of its components) or apreform thereof. The additive manufacturing apparatus 142 may deposit asecond layer 144B of powder within the build space 140 over the firstlayer 144A of at least partially solidified powder. The additivemanufacturing apparatus 142 may thereafter selectively solidify a selectportion of the second layer 144B of powder with the previouslysolidified first portion 148A to form a second portion 148B (e.g.,layer, slice) of the stationary engine structure 34 (or one or more ofits components) or a preform thereof. This process may be repeated untilthe entire stationary engine structure 34 (or one or more of itscomponents) or a preform thereof is formed.

While the additive manufacturing apparatus 142 is described above as alaser powder bed fusion (LPBF) apparatus, the present disclosure is notlimited thereto. The additive manufacturing apparatus 142, for example,may alternatively be configured as a stereolithography (SLA) apparatus,a direct selective laser sintering (DSLS) apparatus, an electron beamsintering (EBS) apparatus, an electron beam melting (EBM) apparatus, alaser engineered net shaping (LENS) apparatus, a laser net shapemanufacturing (LNSM) apparatus, a direct metal deposition (DMD)apparatus or a direct metal laser sintering (DMLS) apparatus.

FIG. 7 is a flow diagram of a method 700 for manufacturing an enginesuch as, but not limited to, the gas turbine engine 20 (or one or moreof its components) of FIG. 1 . This method 700 may be performed using amanufacturing system such as, but not limited to, the manufacturingsystem 136 of FIG. 6 .

In step 702, a stationary engine structure preform 150 is formed. Anexample of the stationary engine structure preform 150 is schematicallyshown in FIGS. 8A and 8B. This stationary engine structure preform 150includes a plurality of engine structure section preforms such as, forexample, an upstream section preform 152, an inner downstream sectionpreform 153 and an outer downstream section 154 preform.

Herein, the term preform may describe a body having at least a basicconfiguration (e.g., shape, size, features, etc.) of a part to beformed. For example, the upstream section preform 152 may generally havethe same configuration as the upstream section 120 to be formed (seeFIGS. 1 and 3 ). The inner downstream section preform 153 may generallyhave the same configuration as the inner downstream section 121 to beformed (see FIGS. 1 and 3 ). The outer downstream section preform 154may generally have the same configuration as the outer downstreamsection 122 to be formed (see FIGS. 3 and 5 ). However, one or more ofthese engine structure section preforms 152-154 may each include one ormore build features, one or more unfinished surfaces, etc. The buildfeatures (e.g., webs, ribs, spares, etc.) may be included in order tosupport the respective preform during the formation step 702. Thesebuild features, however, are removed (e.g., via machining) after theformation step 702 to provide the respective engine structure section120, 121, 122. In addition or alternatively, one or more surfaces,features (e.g., apertures, bosses, mounting features), etc. of theengine structure section preform may be machined (e.g., refined,smoothed, drilled, etc.) to provide the respective engine structuresection 120, 121, 122. Generally speaking, no additional material willneed to be added to engine structure section preforms 120-122 to providethe respective engine structure sections 120-122 following the formationstep 702.

The stationary engine structure preform 150 may be completely formedwithin the build space 140 of the additive manufacturing apparatus 142.The entire stationary engine structure preform 150 including its sectionpreforms 152-154, for example, may be additively manufacturedconcurrently using a layer-by-layer method within the build space 140.Referring to FIG. 8A, the engine structure section preforms 152-154 maybe formed as discrete (e.g., separate, distinct, unattached, etc.)bodies. Each engine structure section preform 152, 153, 154, forexample, may be separated from each adjacent engine structure sectionpreform 152, 153, 154 by a gap. Alternatively, referring to FIG. 8B, oneor more or all of the engine structure section preforms 152-154 may eachbe formed integral with and thereby connected to one or more or all ofthe other engine structure section preforms 152-154. Each enginestructure section preform 152, 153, 154 of FIG. 8B, for example, isconnected to the adjacent engine structure section preform(s) 152, 153,154 by an intermediate structure 156A, 156B (generally referred to as156); e.g., a tie, a support, a bridge, etc.

While the section preforms 152-154 of the stationary engine structurepreform 150 are described above as being formed concurrently within thebuild space 140, the present disclosure is not limited thereto. Forexample, in other embodiments, each of the section preforms 152-154 (orselect ones of the preforms) may be formed within the build space 140separately from another one or more of the section preforms 152-154.Furthermore, while the formation step 702 is described above in relationto additively manufacturing, the present disclosure is not limitedthereto. For example, in other embodiments, one or more or each of thesection preforms 152-154 may also or alternatively be formed usingcasting, machining and/or various other manufacturing techniques.

In step 704, the stationary engine structure 34 and its engine structuresections 120-122 are provided. For example, the various preforms areremoved from the build space 140 of the manufacturing system 136. Inaddition (before or after being removed from the build space 140), oneor more of the following operations may be performed:

-   -   One or more or each of the engine structure section preforms        152-154 may be de-powdered. For example, any remaining,        un-solidified powder trapped with a respective preform may be        removed (e.g., evacuated) from that preform.    -   The engine structure section preforms 152-154 may be separated        from one another where, for example, the preforms are connected        together as shown in FIG. 8B. For example, the intermediate        structures 156A and/or 156B of FIG. 8B may be cut off and/or        ground down to separate the engine structure section preforms.    -   Any build features may be removed from one or more or each of        the engine structure section preforms 152-154.    -   One or more surfaces, one or more features, etc. of one or more        of the engine structure section preforms 152-154 may be        machined, surface treated, heat treated, etc.    -   One or more surfaces of one or more of the engine structure        section preforms 152-154 may be cleaned.    -   One or more surfaces of one or more of the engine structure        section preforms 152-154 may be coated.        By performing one or more of the above operations and/or other        (e.g., finishing) operations, the engine structure section        preforms 152-154 may be turned into their respective engine        structure sections 120-122.

In step 706, the rotating assembly of the elements 42, 44 and 46 and thebearings 48 are installed with one or more of the engine structuresections 120 and/or 121.

In step 708, the engine structure sections 120-122 are mated andconnected to provide the stationary engine structure 34 and, moregenerally, the gas turbine engine 20. The upstream section 120, forexample, is axially abutted against and attached (e.g., mechanicallyfastened) to the inner downstream section 121. The inner downstreamsection 121 is nested within the outer downstream section 122, and theouter downstream section 122 is attached (e.g., bonded) to the innerdownstream section 121.

The gas turbine engine 20 is described above as a single spool,radial-flow turbojet turbine engine for ease of description. The presentdisclosure, however, is not limited to such an exemplary gas turbineengine. The gas turbine engine 20, for example, may alternatively beconfigured as an axial flow gas turbine engine. The gas turbine engine20 may be configured as a direct drive gas turbine engine. The gasturbine engine 20 may alternatively include a gear train that connectsone or more rotors together such that the rotors rotate at differentspeeds. The gas turbine engine 20 may be configured with a single spool(e.g., see FIG. 1 ), two spools, or with more than two spools. The gasturbine engine 20 may be configured as a turbofan engine, a turbojetengine, a propfan engine, a pusher fan engine or any other type ofturbine engine. In addition, while the gas turbine engine 20 isdescribed above with an exemplary reverser flow annular combustor, thegas turbine engine 20 may also or alternatively include any othertype/configuration of annular, tubular (e.g., CAN), axial flow and/orreverser flow combustor. The present disclosure therefore is not limitedto any particular types or configurations of turbine engines.

While various embodiments of the present disclosure have been described,it will be apparent to those of ordinary skill in the art that many moreembodiments and implementations are possible within the scope of thedisclosure. For example, the present disclosure as described hereinincludes several aspects and embodiments that include particularfeatures. Although these features may be described individually, it iswithin the scope of the present disclosure that some or all of thesefeatures may be combined with any one of the aspects and remain withinthe scope of the disclosure. Accordingly, the present disclosure is notto be restricted except in light of the attached claims and theirequivalents.

1. An assembly for a gas turbine engine, comprising: a stationary enginestructure including a diffuser, a combustor, an engine case and aplenum, the combustor disposed within the plenum, the engine caseforming a peripheral boundary of the plenum, and a gas path extendingsequentially through the diffuser, the plenum and the combustor; a firstsection of the stationary engine structure formed as a first monolithicbody, the first section including the diffuser and the combustor; and asecond section of the stationary structure formed as a second monolithicbody, the second section comprising the engine case; wherein the secondsection circumscribes the first section.
 2. The assembly of claim 1,wherein the diffuser includes a first wall, a second wall and aplurality of vanes; and each of the plurality of vanes is within the gaspath and extends between the first wall and the second wall.
 3. Theassembly of claim 1, wherein the combustor is configured as areverse-flow combustor.
 4. The assembly of claim 1, wherein thestationary engine structure further includes a turbine nozzle downstreamof the combustor along the gas path; and the first section furtherincludes the turbine nozzle.
 5. The assembly of claim 4, wherein thediffuser and the turbine nozzle share a common wall.
 6. The assembly ofclaim 4, wherein the turbine nozzle includes a first wall, a second walland a plurality of vanes; and each of the plurality of vanes is withinthe gas path and extends between the first wall and the second wall. 7.The assembly of claim 1, further comprising: a turbine rotor; thestationary engine structure further including a turbine case housing theturbine rotor and forming a peripheral boundary of the gas path; and thefirst section further including the turbine case.
 8. The assembly ofclaim 1, wherein the stationary engine structure further includes anexhaust duct forming a peripheral boundary of the gas path; and thefirst section further including the exhaust duct.
 9. (canceled)
 10. Theassembly of claim 1, wherein the second section is bonded to the firstsection through a butt joint.
 11. The assembly of claim 1, wherein thesecond section is bonded to the first section through a splice joint.12. The assembly of claim 1, wherein the engine case extends axiallyalong and circumferentially about an axis, and the engine case includesa side wall and an end wall; the side wall projects axially out from theend wall to an axial end of the engine case, and the engine case isattached to the first section at the axial end; and the end wallprojects radially in from the side wall to a radial end of the enginecase, and the engine case is attached to the first section at the radialend.
 13. An assembly for a gas turbine engine, comprising: a stationaryengine structure including a diffuser, a combustor, an engine case and aplenum, the combustor disposed within the plenum, the engine caseforming a peripheral boundary of the plenum, and a gas path extendingsequentially through the diffuser, the plenum and the combustor; a firstsection of the stationary engine structure formed as a first monolithicbody, the first section including the diffuser and the combustor; and asecond section of the stationary structure formed as a second monolithicbody, the second section comprising the engine case; wherein thestationary engine structure further includes a fuel conduit and anozzle; wherein the nozzle is configured to receive fuel from the fuelconduit and inject the fuel into a volume of the combustor; and whereinthe second section further includes at least one of the fuel conduit orthe nozzle.
 14. The assembly of claim 13, wherein the stationarystructure further includes a fuel manifold outside of the engine case;the fuel manifold is configured to supply the fuel to the fuel conduit;and the second section further includes the fuel manifold.
 15. Theassembly of claim 13, wherein at least a portion of the fuel conduitprojects out from the engine case into the plenum towards the fuelinjector.
 16. An assembly for a gas turbine engine, comprising: astationary engine structure including a diffuser, a combustor, an enginecase and a plenum, the combustor disposed within the plenum, the enginecase forming a peripheral boundary of the plenum, and a gas pathextending sequentially through the diffuser, the plenum and thecombustor; a first section of the stationary engine structure formed asa first monolithic body, the first section including the diffuser andthe combustor; and a second section of the stationary structure formedas a second monolithic body, the second section comprising the enginecase; wherein the stationary engine structure further includes an inletsection and a compressor case; wherein the compressor case forms aperipheral boundary of the gas path between the inlet section and thediffuser; and wherein a third section of the stationary structureincludes the inlet section and the compressor case, and the thirdsection is attached to the first section.
 17. (canceled)
 18. An assemblyfor a gas turbine engine, comprising: a diffuser including an innerdiffuser wall and an outer diffuser wall; a combustor including an innercombustor wall, an outer combustor wall and a bulkhead extending betweenand connected to the inner combustor wall and the outer combustor wall;a duct wall; an engine wall including a side wall and an end wall, theside wall projecting out from the end wall to the outer diffuser wall,the side wall brazed to the outer diffuser wall, the end wall projectingout from the side wall to the duct wall, and the end wall brazed to theduct wall; and a monolithic body including the diffuser, the combustorand the duct wall.
 19. (canceled)
 20. (canceled)